In a turbine engine, compressors and turbines typically have axially arranged and alternate sets or stages of rotor blades and stator vanes. The stator vanes are mounted to a casing and the rotor blades are mounted to rotor discs. The rotor blades and stator vanes each comprise aerofoils mounted on platforms and the surfaces of which define a working gas flow passage.
The efficiency of the engine is strongly influenced by the shape and the configuration of the aerodynamic surfaces of the rotor blades and stator vanes. The behaviour of the main working gas flow through the compressor and turbine is highly complex and can vary dependent on the engine output, the input of secondary gas flows to the main working gas flow and locally throughout the gas flow passage.
For turbines in particular, additional complexity in the working gas flow can arise from the temperature traverse of the working gas flow from the combustor and thermal characteristics of the turbine blades and stator vanes. Numerous attempts have been made to optimise certain aspects of blade and vane designs to improve stage efficiency and thermal management of the gas flow passage surfaces.
WO0061918A2 discloses a vortex elimination device disposed at the intersection of a blade or vane and its endwall or platform. The vortex elimination device has a generally triangular shape with a straight or curvilinear leading edge and is integral with or attached to the airfoil and endwall. The vortex elimination device prevents the formation of a leading edge vortex as the flow stream passes over the leading edge of the airfoil by generating a radial leading edge force that counters the radial equilibrium and stagnated flow forces, thereby providing a smooth flow stream around the airfoil leading edge.
EP1074697 A2 discloses a method for inhibiting radial transfer of core gas flow away from a center radial region and toward the inner and outer radial boundaries of a core gas flow path. A flow directing structure includes an airfoil having a fillet which diverts the core gas flow away from the area where the airfoil abuts the end wall. Increasing the velocity of the core gas flow in the area where the leading edge of the airfoil abuts the wall impedes the formation of a pressure gradient along the surface of the airfoil that forces core gas from the center region of the core gas toward the wall.
In “Turbine Blade Aerodynamics”, by Sumanta Acharya & Gazi Mahmood, Louisiana State University, CEBA 1419B, Mechanical Engineering Department, pages 363-390 there is disclosed a Leading Edge Fillet or leading edge contouring near the endwall. Fillets are placed at the junction of the leading edge and endwall. Two types of basic construction of fillet profiles can be identified: (i) profile with varying height from the blade surface to the endwall and (ii) profile of bulb with surface thickness at the outer periphery.
However, none of these documents address the problems associated with from the interaction of the main working gas flow and secondary or leakage flow egressing immediately upstream of a set of rotor blades or stator vanes.